1. Field of the Invention
The present invention relates to improvements in radial pulse rocket motors having a plurality of solid propellant units therein and incorporating a thermal barrier that enables the ignition of the propellant units to be independent of each other whereby discrete pulses, specifically "boost" and "sustain" pulses are available upon command. The invention relates particularly to an improvement in the igniter for the sustain phase of such motors.
2. Description of the Prior Art
The entire propulsive capacity of a solid propellant rocket motor, is usually spent during the combustion process of one mass of solid propellant grain. This is for the reason that, once a solid propellant grain is ignited, it is very difficult to stop the combustion process until the entire mass of ignited propellant has been consumed.
Burning may be started by standard initiation means, knonw in the art, and may be of the "end" burn type or the "raidal" burn type. End burning is the burning of a tube or rod of propellant grain in a direction parallel to the axis of the tube or rod, and in a direction away from the nozzle of the rocket motor. Radial burning is the radial or outward burning internally of a tube of propellant grain with the total inner surface thereof being ignited.
It has been proposed in the prior art to provide solid propellant rocket motors of both the end burn and radial burn types with an ability to fire more than once, that is, a rocket motor with a "start-stop-restart" capability by providing two or more concentric layers or zones of solid propellant grain in a combustion chamber with a flame inhibiting or thermal barrier separating the layers, the barrier being made of a material that will confine the burning to a single layer or zone but which is rupturable and destructible so that the next adjacent layer may be ignited with the ruptured and destroyed barrier being blown out of the motor nozzle.
Radial burn prior art rocket motor arrangements that may be fired more than once are disclosed in U.S. Pat. No. 3,248,875 granted on May 3, 1966 to R. D. Wolcott, U.S. Pat. No. 3,293,855 granted on Dec. 27, 1966 to W. E. Cuttill et al, and U.S. Pat. No. 4,357,795 granted on Nov. 9, 1982 to T. W. Bastian et al. In the Wolcott patent heat insulating metal foil and individually associated electrically activated igniter bands are provided between each of several concentric layers of solid propellant grain for igniting, upon command, and in turn, each of the next adjacent layers. Similarly, the Cuttill et al patent discloses a pyrotechnic and an electrically ignitable film between each of several concentric layers of solid propellant for igniting each of the next adjacent layers, in turn. In the Bastian et al patent, a burn inhibitor layer is provided between each of several tandemly positioned layers of solid propellant grain and a separate igniter is provided for "radial" burn of each of the layers in turn. The Bastian patent also discloses the use of a burn inhibitor layer between each of several concentric layers of solid propellant grain and a separate igniter for each layer at the aft end of the motor for "end" burn of each of the next adjacent layers beginning with a central core.
Other similar prior art arrangements are disclosed in U.S. Pat. No. 3,564,845 granted to I. H. Friedman, Jr. et al on Feb. 23, 1971 and U.S. Pat. No. 3,568,448 granted to G. E. Webb, Jr. on Mar. 9, 1971 wherein the inner one of two solid propellant concentric layers that are separated by a flame inhibiting barrier is ignited by an aft end igniter. A rupturable membrane seal and perforated support member assembly is provided to isolate a gas generator from the motor combustion chamber during burning of the inner layer. Activation of a head or forward end igniter activates the gas generator. Presure of the gas produced by the generator ruptures the membrane seal, destroys the flame inhibiting barrier, and causes ignition of the second or outer layer of propellant.
In each of the foregoing prior art patents, the flame inhibiting or thermal barrier is destroyed by being ruptured and blown out of the rocket motor nozzle to enable ignition of the next adjacent propellant layer.
U.S. Pat. No. 3,340,691 granted on Sept. 12, 1697 to G. F. Magnum and U.S. Pat. No. 3,354,647 granted on Nov. 28, 1967 to W. C. Aycock disclose similar arrangements but involve the admission of liquid fuel to the combustion chamber for the destruction of the flame inhibiting barrier and the ignition of the adjacent layer of propellant.
U.S. Pat. No. 3,888,079 granted on Jun. 10, 1975 to W. H. Diesinger discloses a rocket motor having two coaxial combustion chambers, tandemly arranged and each containing a solid propellant charge. A partition or bulkhead is positioned between the two chambers. The partition has closure means for preventing ignition of the second propellant charge during ignition of the first propellant charge and for allowing through flow due to pressure generated by ignition of the second propellant charge.
U.S. application Ser. No. 498,603, filed on May 26, 1983 by M. Fling et al and assigned to the assignee of the present invention, discloses a solid propellant rocket motor which incorporates a plurality of concentrically or tandemly fabricated combustion chambers that are separated by a perforated bulkhead, with each of the chambers containing an individually associated propellant charge and igniter. A pressure responsive membrane positioned in a first one of the chambers closes the perforations in the bulkhead to preclude communication between the chambers when a substantially higher pressure is present in the first one of the chambers than in the second chamber, and allows communication between the chambers when the pressures in the chambers is reversed.
The prior art restartable radial pulse rocket motors utilizing multiple ignition and propellant systems that can be ignited at different intervals to provide more than one pulse, for example, a "boost" and a "sustain" pulse, have a number of problems. On the whole, they are overly complex and require complicated mechanisms and structures that do not interface readily with state-of-the art arm-fire devices, that add undesirably to the weight, bulk or cost of fabrication and installation or assembly and/or detract from the reliability thereof under the low temperatures and vacuum conditions encountered in outer space, particularly after long periods of storage, and that are subject to the possibility of damage resulting from the rupture and destruction of the flame inhibiting or thermal barrier and the subsequent blowing out thereof through the nozzle of the rocket motor.